A purpose of an engine fuel control system is to provide an engine with fuel in a form suitable for combustion and to control the rate of fuel flow for accurate control of engine speed and acceleration. It is known to control the thrust of a gas turbine engine using an Electronic Engine Control (EEC), the thrust of the engine being indirectly measured using shaft speed, Engine Pressure Ratio (EPR) or Turbine Power Ratio (TPR). The EEC also controls (i) the shaft speeds within safe operational limits, and (ii) the temperature and pressure at different parts of the engine to avoid undesirable conditions such as surge or stall, and to ensure the integrity of the engine. Environmental considerations as well as growing power demands of modern aircraft require control systems that are robust and optimised to the operating conditions of the aircraft.
Electronic closed-loop fuel control systems have an integrating action which helps to ensure accurate control of the engine while meeting the pilot's demands for thrust and complying with safety limits. Such systems offer distinct advantages in the achievement of accurate Ndot control under normal operating conditions.
U.S. Pat. No. 5,083,277, which is hereby incorporated by reference, discloses an engine fuel control system in which fuel flow to the engine is controlled by a fuel flow metering valve in response to an aggregate fuel flow demand signal. This signal comprises an element computed in accordance with instantaneous engine speed and an overfuelling element computed in accordance with a pilot's thrust or speed demand.
FIG. 1 shows in more detail an engine fuel control system of the type described in U.S. Pat. No. 5,083,277.
The system of FIG. 1 employs a selector control system, generally indicated at 10 with two competing control loops. The first includes the pilot's engine speed demand lever and signal generator 6 (which provides a demanded high pressure shaft speed signal NHD), engine shaft speed error circuit 30, and gain 62a. The second loop, constituting a shaft acceleration controller, comprises an acceleration limiter loop comparator 46, and an integrator 62b. The high pressure shaft speed signal NH is fedback from the engine to the speed governor loop comparator 30 directly and to the acceleration limiter loop comparator 46 via a differentiator (not shown). Selection of one or the other of these control loops is made on the basis of the lowest fuel flow requirement wins by logic block 42 to provide an overfuelling requirement ΔF. Gain 62a and integrator 62b have respective variable gains Kg and Ka, which are used to map from loop error to ΔF.
For the sake of brevity, a deceleration limiter control loop has been omitted from the drawings of FIG. 1. In practice, a deceleration limiter loop would be very much like the acceleration limiter loop, but would use a negative reference signal, and its output would be compared with the output of the NH governor loop on the basis of the highest fuel flow requirement wins. The result of that comparison being carried forward to the lowest wins logic 42.
An estimated engine steady state fuel flow requirement Fss* is computed by engine model block 24 of the feedback loop generally indicated at 14, and this signal is arithmetically summed to the overfuelling requirement ΔF selected by logic 42 at summing junction 26. The aggregate fuel flow demand signal is supplied to one input of a lowest wins logic block 28 for comparison with a maximum fuel flow signal FLim. A further minimum fuel flow limit signal is compared with the output of block 28 by a highest wins comparison but this is omitted from FIG. 1. The resulting trimmed aggregate fuel flow demand signal Fd is connected to control operation of the fuel system 8 which regulates the flow of fuel to engine 2.
Steady state fuel flow control of the engine is provided by the estimate of the engine steady state fuel flow requirement Fss* against the chosen spool speed, i.e. NH, computed by feedback loop 14. The input to feedback loop 14 is the trimmed aggregate fuel flow demand signal Fd. At summing junction 16 the loop output signal, i.e. an estimated steady state fuel flow signal Fss*, is subtracted from Fd. The difference is the modelled overfuelling demand ΔF*, including any limitations imposed by external factors at e.g. logic circuit 28. ΔF* is multiplied at 18 by the rate of change of engine speed with fuel flow increment NHdot/ΔF to provide an estimate of engine acceleration Nhdot*.
The estimated acceleration NHdot* is integrated by integrator 22 to obtain an estimated engine shaft speed NH* from which the estimated steady state fuel flow requirement Fss* is computed by engine model block 24. Additional signals P1, T1 (inlet pressure and temperature) are shown as inputs to block 24. These signals modify the formula used to compute the Fss* versus NH* characteristic or, alternatively, to select the most appropriate member of a family of such characteristics in accordance with prevailing conditions represented by the inputs.
The computed relationship provided by block 24 is matched as accurately as possible to actual engine behaviour. Thus, there may be further inputs to the block for variable parameters affecting engine performance such as mechanical power offtake, compressor air bleed level, inlet guide vane angle etc., the effect of which must be mirrored in the computed relationship.
Although engine fuel control systems of the type described in U.S. Pat. No. 5,083,277 enable the use of a simple fuel metering valve and can provide accurate control of engine speed and acceleration, under some circumstances, particularly when applied to high bypass ratio turbofans, inadequate control loop stability, in particular phase margins, can result for certain loops when these are tuned to fulfil desired bandwidth requirements.
As well as engine fuel control systems, other types of control system are found in gas turbine engines. For example, engine-casing cooling air control systems allow the clearance between the casing and the blade tips of a turbine of the engine to be adjusted, thereby reducing leakage of working fluid between the casing and the tips, and improving engine efficiency.
Substantial performance degradation is caused by the leakage of working fluid between the rotor and stator assemblies of an engine. A significant part of the leakage is attributed to the radial clearance between the casing and the blades tips of a turbine, generally known as turbine tip clearance. However, while it is desired to reduce turbine tip clearance, some clearance is required to accommodate differential thermal expansion between the rotor and stator assemblies.
At cross-sections along the axis of a turbine, the turbine casing forms a circle surrounding the tips of rotor blades. The blades are in intimate contact with the processed fluid and respond rapidly to variations in working medium temperature. The casing is located at the boundary of the gas path and responds more slowly to changes in operating conditions. The casing is, conventionally, segmented to prevent the build up of mechanical stresses within the turbine casing. As the engine is accelerated during operation, the rotor assemblies grow radially outward toward the casing. A substantial initial clearance is provided between the turbine seal formed on the radially inner wall of the casing and the blade tips to permit this radially outward growth and to allow turbine operation to be free of rub. The minimum clearance between the blade tips and the casing generally occurs during transient operating phases, such as takeoff and acceleration. At equilibrium conditions, however, the clearance tends to increase and excessive leakage of working fluid can occur through the gaps between rotor and stator assemblies.
The radial clearance between the rotor and stator assemblies at cruise conditions can be reduced through the application of turbine casing cooling systems. As the casing is cooled it shrinks, thereby reducing tip clearance to a more acceptable level.
U.S. Pat. No. 5,048,288 discloses a turbine tip clearance control arrangement in which the outer air seal of a gas turbine engine is continuously cooled by compressor discharge air and by a by-pass line which can be throttled or shut off. US patent application no. 2004/0240988 discloses a valve assembly for a turbofan jet for supplying cooling air to turbine case sections.
U.S. Pat. No. 6,487,491 proposes closed-loop tip clearance control, in which case the tip clearance of the engine is monitored during flight, including take off and landing, and adjusted by controlling air flow adjacent the engine casing in response to thermal growth.
Although turbine casing cooling holds considerable promise for improved engine performance, that promise has not yet been fully realised in practice.